This invention relates to the field of electrical testing of remotely sensed liquid containment vessels as are used in aircraft and other vehicles.
In the modern aircraft it is common practice to use the main thrust generating engine as a source of mechanical prime mover energy for driving a plurality of systems necessary for operation of the aircraft. According to this practice the main propulsion engine of the aircraft usually drives a source of electrical energy such as a generator or alternator, one or more hydraulic pumps, an environmental control system (including air conditioning and crew compartment pressurization apparatus) and possibly such additional items as fuel pumps, air compressors, and vacuum pumps.
Such drive arrangements, although quite satisfactory during normal operation of the aircraft, are attended by an inherent difficulty during failure of the aircraft engine even for a temporary time interval. Such failure can of course place the aircraft and its crew in an instantly threatening situation. This occurs particularly in the case of single engined military fighter aircraft where combat or malfunction related loss of main engine would, without other arrangements, make the aircraft instantly uncontrollable and also incapable of crew life support.
Since the failure of hydraulically powered flight control systems and the electrical system of an aircraft is of such critical importance, it is common practice in most aircraft designs to provide for temporary sources of at least electrical and hydraulic system energization--i.e., to provide sources which are not dependent on the main engine of the aircraft. In many World War II vintage aircraft for example it was common practice to include a small gasoline engine energized DC generator within the aircraft, in addition to the normal aircraft storage battery, in order to provide for emergency engine starting and for a continuous source of electrical energy.
In more modern aircraft of the commercial transport variety it has become common practice to include a gas turbine driven alternator or generator for use in both emergency flight conditions and as a source of grounded aircraft energy for cabin air conditioning, engine starting and other functions. In the modern fighter aircraft such as the F-16, it is also desirable to provide an emergency source of electrical energy and hydraulic pressure for use in controlling the aircraft during emergency flight situations--and to operate this source of emergency energy from a fuel supply that is separate and different from that of the main propulsion engine. A system of this type is often made to be automatic in operation in response to engine failure and provides the pilot with at least the capability to control the airborne craft for attitude righting, crew member ejection and possibly for even an emergency landing.
The use of a separate fuel supply for such an emergency power unit is desirable both from the consideration of possible interruption of the main engine fuel and in order to minimize the volume and weight taxes the emergency energy system. In the F-16 aircraft, for example, a tank or a vessel of some 102 pounds weight and 6.5 gallon capacity is provided for this emergency power unit (or EPU) fuel supply and this tank is filled with a fuel of high energy density such as the hydrazine compound that has been used in space vehicles. In the F-16 aircraft this tank or vessel is located in an EPU compartment situated behind the aircraft cockpit.
In view of the corrosive nature, flammability, health, and other hazards attending fuels such as hydrazine, it is found convenient to replenish the supply of these fuels by removing the containment vessel from the aircraft and performing both a fuel replenishment and other restoring of the vessel in a maintenance shop or some other non flight-line environment. Such restoring in a specialized environment is also desirable in view of the relatively small size of this vessel on the F-16 aircraft and especially in view of the need for more than simple fuel replenishment in order to reuse the vessel on the same or another aircraft. In particular in the F-16 aircraft the hydrazine emergency power unit fuel vessel is operated in a sealed condition during normal use of the aircraft and is gas pressurized automatically upon the occurrence of an event requiring emergency power unit operation. This transfer from sealed to pressurized fluid operating condition is achieved by way of an expendable "burst disc" which is ruptured by a charge of pressurized nitrogen gas supplied automatically to the fuel vessel when emergency power unit operation is needed.
For both keeping the aircraft pilot appraised of the possible emergency power unit operating time remaining and for convenient ground check out of the emergency power unit fuel supply it is desirable to indicate the quantity of fuel remaining in this tank to the aircraft pilot by way of a cockpit received gauge or instrument. In the case of hydrazine fuel this remote sensing function can be conveniently provided by way of an electrical capacitance operated probe member received in the emergency power unit "EPU" fuel supply vessel.
Practical experience has shown that normal operation and maintenance of a group of aircraft provided with this EPU system can be attended by a number of inconveniences and labor increasing factors that are deserving of improvement, however. For example, whenever an F-16 aircraft has incurred an EPU automatic start-up event it is necessary for the ground crew of the aircraft to remove the hydrazine tank, replace the pressure sealing "burst disc" in the tank (a step which includes tipping the tank on end to remove hydrazine from the disc area), fill the tank with fuel using a weight determination of 100% filling, and then return the tank to the aircraft and verify that the sensor system instrument in the cockpit indicates a 100% full condition.
In the event of a cockpit instrument indication of something other than 100% tank filling, an event which happens with undesirable frequency in real world settings, it is necessary for the ground crew to dismount the filled tank, a tank weighing 102 Pounds, and identify which of several possible sources for this weight to instrument disagreement is in need of correction. In view of the corrosive and water attracting nature of the hydrazine fuel it is found that the sources of this disagreement may include, for example:
1. A defective capacitance sensor element in the hydrazine tank. PA0 2. A defective indicating instrument in the aircraft cockpit. PA0 3. Defective aircraft wiring between the cockpit instrument and the hydrazine tank. PA0 4. Contamination of the tether cable connector with moisture or hydrazine or both. PA0 5. Similar contamination of the tank connector plug with moisture or hydrazine. PA0 6. Absence of a 100% filling of the tank. PA0 7. Other random and less frequently occurring malfunctions.
In view of the number of these possible difficulties and the relatively large number of possible combinations of these difficulties it is found that considerable time and labor can be saved by eliminating as many of them as possible while the tank is in the refurbishing and refueling shop. Such shop oriented check out is especially desirable where the refueled tank is to be carried perhaps hundreds or thousands of miles from a home base shop to an aircraft which has landed, after an EPU actuation event, at some distant air base or airport. The present invention is found to be of great value in reducing these practical difficulties.
The prior patent art discloses a number of fuel indicating and fuel system check out arrangements which are of general background interest with respect to the present invention. Included in these patents are the float assembly checking system of L. J. Jannotta as described in U.S. Pat. No. 4,821,022, the vehicle fuel quantity indicating apparatus of H. Wamamoto as disclosed in U.S. Pat. No. 4,178,802, and the measuring transducer apparatus of G. Schick as disclosed in U.S. Pat. No. 3,742,342. Also included in these prior patents is the fuel level warning system of R. V. Godferay as disclosed in U.S. Pat. No. 3,333,469, the vehicle fuel indicator of J. C. Montero as disclosed in U.S. Pat. No. 1,414,298 and the electrical indicating device of S. F. Cole as disclosed in U.S. Pat. No. 1,304,022. None of these prior patents, however, teaches the arrangement of an aircraft fuel system check out apparatus and its use as Provided in the present invention.